Radially translating fan nozzle nacelle

ABSTRACT

A variable area nozzle system for a turbofan gas turbine engine comprises a fan duct inner wall, a fan duct outer wall and a fan nozzle. The fan duct outer wall is disposed in radially-spaced relation to the fan duct inner wall. The fan nozzle defines at least a portion of the fan duct outer wall and has a nozzle aft edge defining a fan duct throat area relative to the fan duct inner wall. The fan nozzle is configured to move outwardly relative to the longitudinal axis during axial aft translation thereof in order to vary the fan duct throat area.

CROSS-REFERENCE TO RELATED APPLICATIONS

(Not Applicable)

STATEMENT RE: FEDERALLY SPONSORED RESEARCH/DEVELOPMENT

(Not Applicable)

FIELD

The present disclosure relates generally to gas turbine engines and,more particularly, to an engine nacelle having a variable area fannozzle.

BACKGROUND

Aircraft noise pollution is a significant environmental problem forcommunities near airports. Jet engine exhaust accounts for a majority ofthe noise produced by engine-powered aircraft during takeoff. Because itoccurs at a relatively low frequency, jet engine exhaust noise isunfortunately not effectively damped by the atmosphere alone. The priorart includes several attempts at reducing jet engine exhaust noise. Suchattempts are directed at altering the flow characteristics of the engineexhaust which can be comprised of several components.

Bypass turbofan engines typically produce two exhaust stream components.A first component stream is referred to as the primary exhaust flow andis discharged from a core exhaust nozzle after passing through a coreengine. A second component stream passes through an annular fan ductwhich surrounds the core engine. The second component stream, referredto as the fan exhaust flow, exits a fan nozzle collectively defined byan aft edge of the fan nozzle and the fan duct inner wall whichsurrounds the core engine. The fan exhaust stream and the primaryexhaust stream collectively form the thrust that is generated by theengine.

In bypass turbofan engines, the primary exhaust flow throat area at theexhaust nozzle and the fan exhaust flow throat area at the fan nozzleare preferably optimized for specific engine operating condition. Forexample, during takeoff, a relatively high level of thrust is requiredof the engines as compared to lower levels of thrust that are requiredduring cruise flight. Increasing the quantity or mass of airflow throughthe fan duct having a fixed throat area at the fan nozzle results in anincrease in the velocity of the airflow. An increase in the nozzle exitvelocity results in an increase in the amount of noise that is generatedby the nozzle.

For example, if the fan nozzle throat area is configured for duct massairflow at cruise conditions, then the increased mass of airflowassociated with higher thrust levels will result in a higher velocity ofthe airflow through the fan nozzle. Nozzle exit velocities that arehigher than the optimal velocity for a given nozzle exit area result ina generally higher level of exhaust noise. Noise generated by the fannozzle exhaust may be reduced by decreasing the velocity of airflowthrough the fan nozzle. Increasing the fan nozzle exit or throat arearesults in a reduction in the velocity of the exhaust as it exits thefan duct and therefore reduces the level of noise.

Included in the prior art are several approaches to increasing the fannozzle exit area (i.e., throat area) such as during takeoff in order toreduce exhaust noise. One approach includes linearly translating the fannozzle in an aft direction parallel to a longitudinal axis of the enginein order to increase the fan nozzle exit area and thereby reduce thevelocity of the exhaust. Although effective in reducing exhaust noise,the aft-translating approach presents several deficiencies which detractfrom its overall utility. For example, in some prior art engines, theaft-translating approach results in the creation of a slot or openingwhich allows air to exhaust through the cowl wall. Unfortunately, theopening in the cowl wall adds additional cross-sectional area ratherthan enlarging the exhaust nozzle throat.

Furthermore, the creation of the opening results in leakage through theengine nacelle with an associated loss of engine thrust. Additionally,the aft-translating approach requires the use of swiping seals whichpresent a maintenance risk. An additional drawback associated with theaft-translating approach is that an overlap is created between the ductwall and the fan nozzle resulting in a reduction in the surface area ofacoustic treatment in the fan duct. Such acoustic treatment may includesound-absorbing material such as honeycomb placed along the fan ductinner wall to absorb some of the exhaust noise.

Even further, the aft-translating sleeve must be capable of moving arelatively large distance between stowed and deployed positions in orderto provide optimum noise-reduction/engine thrust capability at takeoffin the deployed position and optimal engine efficiency at cruise in thestowed position. For wing-mounted engines, the presence of trailing edgecontrol surfaces such as wing flaps may present clearance problemsbetween the translating sleeve and the control surface considering theamount of travel of the translating sleeve.

Another approach to increasing the fan nozzle exit area as a means toreduce noise generated during high thrust events such as during takeoffis through the use of expanding flaps or petals which form the nozzleexit external surface. More typically applied to primary exhaust nozzlesof military aircraft, the flaps or petals may be pivoted outwardly toenlarge the throat area of the nozzle and thereby reduce the exhaustvelocity. The flaps or petals may also be biased to one side or theother in order to provide thrust vectoring for increased maneuverabilityof the aircraft. As may be appreciated, the implementation of a flap orpetal scheme for changing nozzle exit area is structurally andfunctionally complex and presents weight, maintenance and cost issues.

An additional consideration in a variable area fan nozzle for reducingexhaust noise is that a movable fan nozzle must be compatible withthrust reversers commonly employed on modern jet engines. As is known inthe art, thrust reversers on jet engines may reduce landing distance ofan aircraft in normal (e.g., dry) runway conditions or increase safetyin slowing the aircraft in slick (e.g., wet) runway conditions. Thrustreversers operate by reorienting the normally aftwardly directed flow ofexhaust gasses into a forward direction in order to provide brakingthrust to the aircraft. The reorienting of the engine exhaust gasses isfacilitated by spoiling, deflecting and/or turning the flow stream ofthe primary exhaust and/or the fan exhaust.

For turbofan engines, thrust reversers may include the use of cascades,pivoting doors or by reversing the pitch of the fan blades. Incascade-type thrust reverser, the turbofan engine may include an outertranslating sleeve which is configured to move axially aft to uncoverdeflecting vanes mounted in the nacelle cowl. Simultaneous with the aftmovement of the translating sleeve, blocker doors in the fan duct areclosed in order to redirect the fan flow outwardly through thedeflecting vanes and into a forward direction to providethrust-reversing force. Due to the widespread implementation of thrustreversal capability on many aircraft, a variable area fan nozzle must becompatible with thrust reverser systems commonly employed on modern jetengines

As can be seen, there exists a need in the art for a variable area fannozzle which is effective in increasing the nozzle exit area of a gasturbine engine in order to reduce noise at takeoff by reducing exhaustvelocity. In addition, there exists a need in the art for a variablearea fan nozzle which can achieve an increase in nozzle area but whichrequires a minimal amount of travel to avoid interfering with variouscomponents such as trailing edge control surfaces. Also, there exists aneed in the art for a variable area fan nozzle which is compatible withthrust reversers commonly employed on gas turbine engines. Finally,there exists a need in the art for a variable area fan nozzle which issimple in construction and requiring minimal maintenance.

BRIEF SUMMARY

The above-noted needs associated with fan nozzles of the prior art arespecifically addressed and alleviated by the present disclosure whichprovides a variable area nozzle system for a gas turbine engine whereina fan nozzle of the nozzle system is specifically configured to moveoutwardly with simultaneous aft movement of the fan nozzle in order tovary the fan duct throat area of the gas turbine engine. In this regard,the fan nozzle is configured to move between stowed and deployedpositions.

The fan nozzle is configured to move outwardly relative to alongitudinal axis of a gas turbine engine while the fan nozzle issimultaneously translated axially in an aft direction from the stowedposition to the deployed position in order to vary the fan duct throatarea. Likewise, the fan nozzle is configured to move inwardly duringaxially forward translation back to the stowed position. The fan nozzle,in a preferable embodiment, maintains the same orientation when movedbetween the stowed and deployed positions. In this regard, the fannozzle maintains an angled orientation relative to the longitudinal axisduring the aft-outward movement and forward-inward movement.

The technical effects of the disclosure include a reduction in theamount of movement required in order to achieve a given increase in fannozzle throat area as compared to prior art fan nozzles which are purelyaxially translating (i.e., purely forward and aft motion of the fannozzle). Furthermore, the fan nozzle as disclosed herein provides arelatively simple structural arrangement with relatively few movingparts as compared to more complex variable fan nozzle arrangements whichcomprise moveable petals for varying the fan nozzle exit area.

Advantageously, the fan nozzle is adapted to be extended aftwardly froma translating sleeve (i.e., a cowl assembly) of a thrust reverser for agas turbine engine. In this regard, the fan nozzle is adapted to beoperated independently of the thrust reverser and may be moved from thestowed to deployed positions with the thrust reverser either in the openor closed position.

The nozzle system comprises a fan duct inner wall and a fan duct outerwall disposed in radially-spaced relation to the fan duct inner wall.The nozzle system is adapted such that the fan duct outer wall issubstantially continuous in order to provide smooth aerodynamic flowacross the fan nozzle and the translating sleeve when the fan nozzle ismoved between the stowed and deployed positions. Sealing mechanismsprovided between the fan nozzle at the nozzle forward edge and the cowlassembly (i.e., translating sleeve) on an aft end thereof minimize anyleakage which may otherwise reduce engine efficiency.

In one embodiment, the nozzle system may comprise a first seal which maybe configured as a bulb seal disposable in sealing contact between thenozzle forward edge and a cowl outer panel of the translating sleevesuch that sealing engagement between the fan nozzle and the translatingsleeve is provided when the fan nozzle is moved to the deployedposition.

Likewise, a second seal configured in an optional bulb seal embodimentmay be provided between the fan nozzle and the aft edge of the cowlinner panel in order to provide sealing engagement between the fannozzle and the cowl assembly (i.e., translating sleeve) when the fannozzle is moved to the stowed position. In this manner, the nozzlesystem provides for sealing of the fan nozzle to the cowl assembly ortranslating sleeve in the stowed and/or deployed positions in order tomaintain aerodynamic efficiency and internal fan duct pressure.

Advantageously, the variable area fan nozzle as disclosed herein allowsfor at least two different fan duct throat areas. More specifically,when the fan nozzle is in the stowed position, the fan duct throat areais preferably optimized for cruise flight where noise reduction is notan issue. When the fan nozzle is moved to the deployed position, the fanduct throat area is increased in order to reduce the velocity of the fanflow exhaust out of the fan nozzle to thereby reduce the level of noise.

In this regard, the larger fan duct throat area may be selected forhigher power settings such as for takeoff where noise suppression isrequired. A smaller fan duct throat area may be selected for lower powersettings such as during cruise flight where noise is not an issue butoptimal engine performance dictates a smaller nozzle throat area.

The features, functions and advantages that have been discussed can beachieved independently in various embodiments of the present inventionor may be combined in yet other embodiments, further details of whichcan be seen with reference to the following description and drawingsbelow.

BRIEF DESCRIPTION OF THE DRAWINGS

These and other features of the present invention will become moreapparent upon reference to the drawings wherein like numbers refer tolike parts throughout and wherein:

FIG. 1 is an aft perspective illustration of a turbofan gas turbineengine mounted on a strut or pylori of an aircraft;

FIG. 2 is a side view of the gas turbine engine illustrating a fannozzle and a primary exhaust nozzle disposed aftwardly of the fannozzle;

FIG. 3 is a side view of the gas turbine engine illustrating the fannozzle moved outwardly and aftwardly into its deployed position ascompared to the stowed position of the fan nozzle illustrated in FIG. 2;

FIG. 4 is a sectional illustration taken along lines 4-4 of FIG. 2 andillustrating the fan nozzle extending aftwardly from a fan duct andwherein the fan nozzle is illustrated in the stowed position;

FIG. 4A is an enlarged sectional illustration of a first seal configuredto provide sealing of the fan nozzle in the deployed position;

FIG. 5 is a sectional illustration taken along lines 5-5 of FIG. 3 andillustrating the fan nozzle moved outwardly and aftwardly relative to alongitudinal axis of the engine wherein the fan nozzle is moved to thedeployed position;

FIG. 6 is a sectional illustration of the fan nozzle illustrating athrust reverser in an open position;

FIG. 7 is a sectional illustration taken along lines 7-7 of FIG. 5 andillustrating a slider mechanism in one embodiment comprising a tongueelement and a groove element slidable in relation to one another; and

FIG. 8 is a sectional illustration taken along lines 8-8 of FIG. 3 andillustrating a fan nozzle forward edge extending between a cowl innerpanel and a cowl outer panel.

DETAILED DESCRIPTION

Referring now to the drawings wherein the showings are for purposes ofillustrating preferred and various embodiments of the disclosure onlyand not for purposes of limiting the same, shown in FIG. 1 is a highbypass turbofan gas turbine engine 10 incorporating a variable areanozzle system 82. In a broad sense, the nozzle system 82 includes a fannozzle 84 which is specifically configured to move radially outwardlysimultaneous with axial movement thereof. More specifically, the fannozzle 84 is configured to move radially outwardly during axially afttranslation and radially inwardly during axially forward translationbetween stowed and deployed positions 98, 100.

The fan nozzle 84 defines at least a portion of a fan duct outer wall 42of the turbofan gas turbine engine 10. The fan nozzle 84 includes anozzle aft edge 88 which, together with a fan duct inner wall 40,defines a fan duct throat area Ta that is increased by the radiallyoutward/axially aft translating motion of the fan nozzle 84. Byincreasing the fan duct throat area Ta, the velocity of fan flow 44passing through the fan duct 38 decreases and therefore reduces thelevel of exhaust noise generated thereby.

Although the present disclosure is described in the context of a highbypass gas turbine engine 10 as illustrated in the Figures, the nozzlesystem 82 may be implemented on other types of gas turbine engineswherein it is desirable to increase the fan duct throat area Ta to anoptimal size for a given engine operating condition. For example, thenozzle system 82 may be configured to allow for movement of the fannozzle 84 to a deployed position 100 wherein the throat area Ta of thefan nozzle 84 is at a maximum as may be desirable for high thrustsettings of the engine 10 such as during takeoff and climb-out but wherenoise suppression is also desired in order to reduce environmental noiseimpact. Likewise, the fan nozzle 84 may be moved to the stowed position98 wherein the throat area Ta of the fan nozzle 84 is minimized orplaced in an optimal position for lower engine thrust settings as may bedesirable for cruise flight where noise suppression is not required butwhere nozzle efficiency dictates a reduced throat area of the fan nozzle84.

Referring to FIG. 1, shown is the gas turbine engine 10 supported by apylori or strut 12 which, in turn, may be mounted to an aircraft 118. Asis known in the art, the turbofan gas turbine engine 10 may include acore engine 28 within which pressurized air may be mixed with fuel forgenerating combustion gases which flow through turbine stages and areexpelled at a primary exhaust nozzle 32. As can be seen in FIG. 1, theprimary exhaust nozzle 32 may be defined by a generally conically shapedprimary exhaust plug 30 at an aft end of the fan duct inner wall 40.

The fan duct 38 is bounded by the fan duct inner wall 40 and the fanduct outer wall 42. A fan flow 44 passes through the fan duct 38 throughthe influence of air that is pressurized by a fan (not shown) located ata forward end of the engine 10 near an inlet 22 of the engine 10. Alarge portion of the propulsive thrust generated by the engine 10 is aresult of the pressurized air passing through the fan duct 38 andexiting the fan nozzle 84 which is illustrated with chevrons 92 on thenozzle aft edge 88.

The gas turbine engine 10 includes a nacelle 14 having the inlet 22 atthe forwardmost end of the engine 10 and a stationary fan cowl 36 whichhouses the rotating fan blades (not shown). The stationary fan cowl 36forms part of a cowl assembly 34 which preferably defines at least aportion of the fan duct outer wall 42. The cowl assembly 34 may furtherinclude a translating sleeve 46 for gas turbine engines having thrustreverser capability. The thrust reverser 62 comprising the translatingsleeve 46 is configured to move axially aftwardly in order to redirectthe fan flow 44 through one or more cascade segments 74 comprising aplurality of deflecting vanes mounted in a circumferential manner aboutthe fan duct 38. Activation of the thrust reverser 62 causes the fanflow 44 to be reoriented into a forward direction when the cascadesegments 74 are uncovered by the axially translating sleeve 46. Thethrust reverser 62 including the translating sleeve 46 may be actuatedby one or more thrust reverser actuators 64 shown in FIGS. 1 and 8 onupper, center and lower sections of left and right engine halves.

Although the engine 10 may be configured such that the inner fan duct 38comprises an unobstructed annular geometry, the engine 10 may beprovided in a bifurcated arrangement wherein the fan duct 38 is dividedinto two generally semi-cylindrical annular cavities joined at an upperbifurcation 18 along a hinge beam 24 and at a lower bifurcation 20 alonga latch beam 26. In this manner, the fan cowl 36 and the fan duct 38including the cowl assembly 34 and the translating sleeve 46 may beunlatched at the latch beam 26 on the lower end of the nacelle 14 andmay be pivoted or hinged upwardly in order to expose the core engine 28for maintenance purposes. Each of the latch beam 26 and hinge beams 24may include a cowl slider track 52 and cowl slider beam 54 to which thetranslating sleeve 46 portion of the cowl assembly 34 including the fannozzle 84 may be moved axially aft in order to uncover the cascadesegments 74 during thrust reverser actuation.

Referring briefly to FIG. 6, shown is the thrust reverser 62 in an openposition wherein it can be seen that the cascade segments 74 areuncovered due to aft axial translation of the translating sleeve 46. Thecross-sectional illustration of FIG. 6 illustrates one of the thrustreverser actuators 64 fixedly mounted adjacent a bull nose 70/torque boxof the fan cowl 36 and coupled at an opposite end via a rod end 66coupled to an actuator bracket 68. The actuator bracket 68 may befixedly mounted between a cowl outer panel 56 and a cowl inner panel 58.The cowl outer panel 56 and cowl inner panel 58 collectively define thetranslating sleeve 46.

Upon activation of the thrust reverser 62 such that the translatingsleeve 46 is moved axially aftwardly, one or more blocker doors 76 arepivoted downwardly into contact with the fan duct inner wall 40 due todrag links 78 pivotably connected to the fan duct inner wall 40 at oneend and to the blocker door 76 at an opposite end of the drag link 78.In the deployed position 100, the blocker doors 76 prevent passage offan flow 44 to the aft end of the fan duct 38 and instead cause the fanflow 44 to be redirected outwardly through the cascade segments 74 in aforward motion for thrust reversal purposes.

The cascade segments 74 are supported by a cascade support ring 72 at anaft end of the cascade segments 74 and by the bull nose 70 and torquebox at a forward end of the cascade segments 74. The semi-cylindricaltranslating sleeve 46 and fan nozzle 84 may be supported at upper andlower sides (i.e., along the upper and lower bifurcations) by a cowlslider beam 54 that slides along a cowl slider track 52 disposed at thehinge beam 24 and latch beam 26 on each half of the engine 10.

Referring to FIGS. 2-5, shown are side views and sectional views of thegas turbine engine 10 illustrating the fan nozzle 84 configured to movesimultaneously radially outwardly and axially aftwardly relative to alongitudinal axis 16 of the engine 10. As was earlier indicated, the fannozzle 84 is specifically adapted to move in an angled relation to thelongitudinal axis 16 in an aft-outward and forward-inward direction. Inthis manner, the fan nozzle 84 is adapted to vary the fan duct throatarea Ta. Preferably, the fan nozzle 84 is configured to move betweenstowed and deployed positions 98, 100 while maintaining the sameorientation of the fan nozzle 84 relative to the longitudinal axis 16during the aft-outward movement and during the forward-inward movement.

FIG. 2 illustrates the fan nozzle 84 in the stowed position 98. The fannozzle 84 extends aftwardly from the translating sleeve 46. Thetranslating sleeve 46 includes a sleeve forward edge 48 which isdisposed adjacent to the stationary fan cowl 36. The translating sleeve46 also includes a sleeve aft edge 50 which is disposed in slightlyoverlapping relationship with a nozzle forward edge 86. Referringbriefly to FIG. 4, shown is a cross-sectional illustration of the fannozzle 84 and its interconnectivity with the translating sleeve 46. Ascan be seen, the nozzle forward edge 86 is extendable (e.g., slidable)through an annular slot or opening that may be formed in the translatingsleeve 46.

In one embodiment illustrated in FIGS. 4 and 8, the translating sleeve46 may comprise the cowl outer panel 56 and the cowl inner panel 58.FIG. 8 illustrates the fan nozzle forward edge 86 extending between thecowl outer panel 56 and cowl inner panel 58. The nozzle forward edge 86may be formed as a flange which may be angled slightly inwardly from theexterior nacelle 14 contour defined by the fan nozzle 84 and the cowlouter panel 56. On an interior side of the fan nozzle 84 is a fan nozzlelip 90 which may be disposed in slightly overlapping relationship to aninner panel aft edge 60. The fan nozzle lip 90 may be configured to forma continuation of the fan duct outer wall 42 in order to providecontinuity of fan flow 44 through the fan duct 38. When the fan nozzle84 is in the stowed position 98 as illustrated in FIG. 4, the insidegeometry of the fan nozzle 84 adjacent the nozzle forward edge 86 iscompatible with the geometry of the cowl outer panel 56 and cowl innerpanel 58.

Referring to FIGS. 4 and 4A, shown is a first seal 114 disposed along anoutside surface of the nozzle forward edge 86. The first seal 114 may beconfigured to provide sealing contact with the inside surface of thecowl outer panel 56 when the fan nozzle 84 is moved to the deployedposition 100 shown in FIG. 5. The first seal 114 is specifically adaptedto prevent loss of pressure within the fan duct 38 as well as to provideaerodynamic sealing of flow passing along the external nacelle 14surface. An access cover 80 may be included in the area of the firstseal 114 for inspection and maintenance of the first seal 114. In oneembodiment, the first seal 114 may be configured as a bulb seal whichmay be mounted on the nozzle forward edge 86 or on the cowl outer panel56 in such a manner as to prevent rolling of the bulb seal which mayotherwise increase wear of the first seal 114.

In this regard, the cowl outer panel 56 and/or the nozzle forward edge86 may include appropriate geometry to provide proper sealing of thebulb seal to avoid leakage. The seal is preferably configured to providesealing capability during extreme temperature differentials (i.e., 160°F.) as well as provide adequate compatibility with the flexing nacelle14 structure in response to dynamic and static loads. Furthermore, thebulb seal is preferably configured to adapt to the complex curvedgeometry of the engine nacelle 14 and more specifically, to theinterface between the nozzle forward edge 86 and the cowl outer panel56.

Referring briefly to FIG. 5, shown is the fan nozzle 84 in the deployedposition 100. As can be seen, a second seal 116 may be provided betweenthe aft edge of the cowl inner panel 58 and the fan nozzle 84 in orderto provide sealing between the components when the fan nozzle 84 is inthe stowed position 98 illustrated in FIGS. 4 and 4A. The second seal116 is also preferably configured to maintain sealing engagement betweenthe fan nozzle 84 and the translating sleeve 46 when the fan nozzle 84is in the stowed position 98 in order to avoid loss of engine efficiency(i.e., reduced thrust).

The second seal 116 is also preferably configured and mounted in amanner that prevents rolling motion of the second seal 116 as the secondseal 116 engages with the cowl inner panel 58 and/or with the insidesurface of the fan nozzle 84. Although the first and second seals 114,116 are described as bulb seals, any seal geometry or configuration maybe provided and is not solely limited to that which is illustrated inFIGS. 4, 4A and 5. Furthermore, the first and second seals 114, 116 maybe mounted at any location and are not limited to the positions shown inthe figures.

Referring briefly to FIG. 3, shown is a side view of the engine nacelle14 illustrating the fan nozzle 84 in the deployed position 100 where itcan be seen that the sleeve aft edge 50 and the nozzle forward edge 86are disposed in relatively closely-spaced relationship to one another ascompared to the increased spacing illustrated in FIG. 2. Alsoillustrated in FIG. 3 is a plurality of fan nozzle slider mechanisms 104which may facilitate movement of the fan nozzle 84 relative to thetranslating sleeve 46.

Referring to FIGS. 4, 4A, 5 and 7, shown is a slider mechanism 104 forslidably connecting the fan nozzle 84 to the cowl inner panel 58. Asbest seen in FIG. 7, the slider mechanism 104 may be configured as atongue element 106 (i.e., guide element) which is slidably connected toa groove element 108 (i.e., slot element). The groove element 108 may beconfigured as a bracket mounted to the cowl inner panel 58 while thetongue element 106 may be configured as a bracket mounted to the fannozzle 84 at the nozzle forward edge 86. As can be seen in FIG. 7, thenozzle forward edge 86 may include a localized relief for accommodatingthe nozzle slider mechanism 104. The cowl outer panel 56 may be disposedin abutting slidable contact with the exterior surface of the fan nozzle84 to provide aerodynamic continuity along the exterior surface of thenacelle 14.

Referring to FIGS. 2 and 3, the slider mechanisms 104 are illustrated asbeing installed in one embodiment in angularly-spaced relation to oneanother. The slider mechanisms 104 are preferably sized and configuredto react the relatively large internal pressure forces acting againstthe fan cowl 36 and which are directed generally radially outwardly. Inthis regard, the slider mechanisms 104 are preferably adapted toaccommodate 3 or 4 Pascals of pressure acting against the fan nozzle.

In addition, the slider mechanisms 104 are preferably configured toavoid coupling of pressure frequencies in the fan duct 38 with aresonant frequency of the fan nozzle 84 which would otherwise increasethe loads imposed on the slider mechanisms 104. It should be noted thatalthough the slider mechanisms 104 are illustrated as being disposed inangularly-spaced relation along each fan nozzle 84 forward edge, theslider mechanisms 104 may be disposed at any location capable ofresisting the loads imposed on the fan nozzle. Furthermore, although theslider mechanisms 104 are illustrated as comprising a tongue 106 andgroove element 108, alternative embodiments of the slider mechanisms 104are contemplated.

Referring to FIGS. 4-5, it can be seen that when the fan nozzle 84 ismoved between the stowed and deployed positions 98, 100, the fan ductouter wall 42 is substantially continuous across the fan nozzle 84 andthe translating sleeve 46 (i.e., cowl assembly 34) in order to provideuninterrupted flow through the fan duct 38. It can also be seen that thefan nozzle 84 preferably moves along a direction of travel 94 which maybe parallel to a tangency line 96 (i.e., parallel to the cowl innerpanel 58). In one embodiment, the junction of the fan nozzle 84 with thecowl inner panel 58 is preferably positioned at an area of maximumdiameter D_(Tmax) of the fan duct inner wall 40 as shown in FIG. 5. Asshown, the aft end of the cowl inner panel 58 may be angled outwardlyrelative to the longitudinal axis 16 (i.e., engine centerline) of theengine 10. However, the junction between the fan nozzle 84 and the cowlinner panel 58 may be positioned at any location along the fan ductinner wall 40.

Referring briefly to FIG. 5, shown are the relative positions of the fannozzle 84 in the stowed and deployed positions 98, 100. The fan nozzle84 in the stowed position 98 results in a fan nozzle 84 throat areadesignated by Ta_(C) which may be a desired position of the fan nozzle84 where noise is not an issue such as during cruise flight of theaircraft 118. The fan nozzle 84 in the deployed position 100 results inan increased fan nozzle 84 throat area designated by Ta_(T) which may bea desired position of the fan nozzle 84 where noise reduction is desiredsuch as during takeoff and climb-out of the aircraft 118.

As was earlier indicated, the configuration of the fan nozzle 84 resultsin a relatively large increase in nozzle throat area Ta due to theaft-outward movement of the fan nozzle 84 as compared to prior art fannozzle configurations which are purely axially translating. For example,in one embodiment, the fan nozzle 84 as disclosed herein may provide two(2) to three (3) times the increase in fan nozzle throat area Ta thanthe amount of increase provided by prior art fan nozzle configurationswhich are purely axially translating. Furthermore, depending upon thecurvature of the fan duct inner wall 40, a purely aft translating fannozzle may be unable to achieve the amount of increase in fan nozzlethroat area Ta that may be provided by the fan nozzle 84 of the presentdisclosure.

As illustrated in FIG. 6, the fan nozzle 84 extends aftwardly from thetranslating sleeve 46 and is movable between the stowed and deployedpositions 98, 100 when the translating sleeve 46 is moved axially aftfor thrust reversal purposes. In one embodiment, it is contemplated thatthe thrust reverser actuator 64 may be co-located with the fan nozzleactuator 102. For example, the thrust reverser actuator 64 may include atelescoping connecting member 110 which is extendable outwardly from arod end 66 of the thrust reverser actuator 64 as illustrated in FIGS.4-6.

The connecting member 110 may be connected to the fan nozzle 84 forwardedge by means of a link 112 having pivoting capability to accommodatemovement of the nozzle forward edge 86 between stowed and deployedpositions 98, 100 of the nozzle forward edge 86. However, it should berecognized that the fan nozzle actuator 102 illustrated in FIGS. 4, 5and 6 is exemplary in nature and should not to be construed as limitingother alternative embodiments of the fan nozzle actuator 102 forfacilitating the aft-outward and forward-inward motion of the fan nozzle84. Regardless of the specific actuation mechanism for the fan nozzle 84and thrust reverser 62, it is contemplated that the thrust reverseractuator 64 and fan nozzle actuator 102 are separate systems in order toavoid uncommanded activation of the thrust reverser 62.

In one embodiment, the nozzle system 82 may include three fan nozzleactuators 102 for the fan nozzle 84 on each side (i.e., left and rightsides) of the engine 10 in order to avoid skewed deployment orretraction which may result in binding of the fan nozzle 84. It shouldalso be noted that although the fan nozzle 84 is shown as being in thestowed or deployed positions 98, 100 in FIGS. 4 and 5, the fan nozzle 84may be configured to be selectively movable to a predetermined positioncorresponding to an operating parameter of a gas turbine engine 10. Forexample, the fan nozzle 84 may be moved to any one of a plurality ofintermediate positions between the stowed and deployed positions 98, 100and is not limited to either one or the other extreme. Advantageously,the configuration of the fan nozzle 84 avoids any reduction in acoustictreatment which may be applied along the fan duct inner wall 40.

The present disclosure also provides a methodology for varying the fanduct throat area Ta of a gas turbine engine 10. As was indicated above,the gas turbine engine 10 may include the fan duct inner wall 40 and thefan duct outer wall 42 disposed in radially-spaced relation to the fanduct inner wall 40. The fan nozzle 84 may define at least a portion ofthe fan duct outer wall 42. The method may comprise the step of movingthe fan nozzle 84 outwardly relative to the longitudinal axis 16 duringaxially aft translation of the fan nozzle 84. In this regard, the methodmay comprise moving the fan nozzle 84 in the aft-outward motion and/orin the inward-forward motion.

The method may further comprise the step of maintaining the orientationof the fan nozzle 84 relative to the longitudinal axis 16 when movingthe fan nozzle 84 between stowed and deployed positions 98, 100. In thisregard, the fan nozzle 84 is preferably maintained in the same angledrelationship with the longitudinal axis 16 (i.e., engine centerline)when moved between the stowed and the deployed positions 98, 100. Themethod may further comprise the step of maintaining the continuity ofthe fan duct outer wall 42 across the fan nozzle 84 and the cowlassembly 34 (i.e., translating sleeve 46) when moving the fan nozzle 84between the stowed and deployed positions 98, 100.

The above description is given by way of example and not limitation.Given the above disclosure, one skilled in the art could devisevariations that are within the scope and spirit of the embodimentsdisclosed herein. Furthermore, the various features of the embodimentsdisclosed herein can be used alone or in any varying combinations witheach other and are not intended to be limited to the specificcombinations described herein. Thus, the scope of the claims is not tobe limited by the illustrated embodiments.

What is claimed is:
 1. A nozzle system for a gas turbine engine having alongitudinal axis, the nozzle system comprising: a thrust reverser; afan duct inner wall; a fan duct outer wall disposed in radially spacedrelation to the fan duct inner wall, the fan duct inner and outer wallcollectively defining a fan duct; and a fan nozzle defining at least aportion of the fan duct outer wall and having a nozzle aft edge, the fanduct inner wall and the nozzle aft edge collectively defining a fan ductthroat area, the fan nozzle being configured to non-pivotably moveoutwardly relative to the longitudinal axis during axially afttranslation thereof to vary the fan duct throat area; the fan nozzlebeing movable without activating the thrust reverser, the thrustreverser having movable blocker doors configured to block fan flowthrough the fan duct when the thrust reverser is activated.
 2. Thenozzle system of claim 1 wherein: the fan nozzle is configured to moveoutwardly during axially aft translation and inwardly during axiallyforward translation between stowed and deployed positions.
 3. The nozzlesystem of claim 2 wherein: the fan nozzle is configured to maintain anorientation relative to the longitudinal axis when moved between thestowed and deployed positions.
 4. The nozzle system of claim 2 furthercomprising: a cowl assembly defining at least a portion of the fan ductouter wall; wherein: the fan duct outer wall is substantially continuousacross the fan nozzle and the cowl assembly when the fan nozzle is movedbetween the stowed and deployed positions.
 5. The nozzle system of claim4 further comprising: a fan nozzle slider mechanism slidably connectingthe fan nozzle to the cowl assembly.
 6. The nozzle system of claim 1wherein: the fan nozzle is selectively movable to a predeterminedposition corresponding to an operating parameter of the gas turbineengine.
 7. A nozzle system for a gas turbine engine having alongitudinal axis, the nozzle system comprising: a fan duct inner wall;a fan duct outer wall disposed in radially spaced relation to the fanduct inner wall; a fan nozzle defining at least a portion of the fanduct outer wall and having a nozzle aft edge, the fan duct inner walland the nozzle aft edge collectively defining a fan duct throat area,the fan nozzle being configured to move outwardly relative to thelongitudinal axis during axial translation thereof to vary the fan ductthroat area, the fan nozzle being configured to move outwardly duringaxially aft translation and inwardly during axially forward translationbetween stowed and deployed positions; and a cowl assembly defining atleast a portion of the fan duct outer wall; the fan duct outer wallbeing substantially continuous across the fan nozzle and the cowlassembly when the fan nozzle is moved between the stowed and deployedpositions; the cowl assembly including a translating sleeve configuredto move axially aft for thrust reversal; and the fan nozzle extendingaftwardly from the translating sleeve and being movable between thestowed and deployed positions when the translating sleeve is movedaxially aft.
 8. The nozzle system of claim 7 wherein: the translatingsleeve has a sleeve aft edge defining a tangency line; the fan nozzlebeing configured to move along a direction parallel to the tangencyline.
 9. An aircraft having a gas turbine engine with a longitudinalaxis, comprising: a nozzle system including: a thrust reverser; a fanduct inner wall; a fan duct outer wall disposed in radially spacedrelation to the fan duct inner wall, the fan duct inner and outer wallcollectively defining a fan; and a fan nozzle defining at least aportion of the fan duct outer wall and having a nozzle aft edge defininga fan duct throat area relative to the fan duct inner wall, the fannozzle being configured to non-pivotably move outwardly relative to thelongitudinal axis during axially aft translation of the fan nozzle; thefan nozzle being movable without activating the thrust reverser, thethrust reverser having movable blocker doors configured to block fanflow through the fan duct when the thrust reverser is activated.
 10. Theaircraft of claim 9 wherein: the fan nozzle is configured to moveoutwardly during axially aft translation and inwardly during axiallyforward translation between stowed and deployed positions.
 11. Theaircraft of claim 10 wherein: the fan nozzle is configured to maintainan orientation relative to the longitudinal axis when moved between thestowed and deployed positions.
 12. The aircraft of claim 9 furthercomprising: a cowl assembly defining at least a portion of the fan ductouter wall; wherein: the fan duct outer wall of the fan nozzle and cowlassembly are substantially continuous when the fan nozzle is movedbetween the stowed and deployed positions.
 13. The aircraft of claim 12further comprising: a fan nozzle slider mechanism slidably mounting thefan nozzle to the cowl assembly.
 14. The aircraft of claim 9 wherein:the fan nozzle is selectively movable to a predetermined positioncorresponding to an operating parameter of the gas turbine engine. 15.An aircraft having a gas turbine engine with a longitudinal axis,comprising: a nozzle system including: a fan duct inner wall; a fan ductouter wall disposed in radially spaced relation to the fan duct innerwall; a fan nozzle defining at least a portion of the fan duct outerwall and having a nozzle aft edge, the fan duct inner wall and thenozzle aft edge collectively defining a fan duct throat area, the fannozzle being configured to move outwardly relative to the longitudinalaxis during axial translation thereof to vary the fan duct throat area,the fan nozzle being configured to move outwardly during axially afttranslation and inwardly during axially forward translation betweenstowed and deployed positions; and a cowl assembly defining at least aportion of the fan duct outer wall; the fan duct outer wall beingsubstantially continuous across the fan nozzle and the cowl assemblywhen the fan nozzle is moved between the stowed and deployed positions;the cowl assembly including a translating sleeve configured to moveaxially aft for thrust reversal; and the fan nozzle extending aftwardlyfrom the translating sleeve and being movable between the stowed anddeployed positions when the translating sleeve is moved axially aft. 16.The aircraft of claim 15 wherein: the translating sleeve having a sleeveaft edge defining a tangency line; the fan nozzle being configured tomove along a direction parallel to the tangency line.